A thruster is a device which energizes a propellant such that when the propellant is ejected from the thruster, momentum is generated to move the body to which the thruster is attached. Thrusters use many different kinds of mechanisms to energize the propellant, but one common type of thruster introduces an electric current to the propellant to energize the propellant. These electric thrusters are commonly used in man-made satellites.
Electric thrusters can generally be categorized into two groups: steady state thrusters and pulsed thrusters, Each has its advantages and disadvantages.
As the name suggests, a steady state thruster is a thruster wherein the propellant is energized by providing a steady state electrical current to the propellant. One such steady state thruster is shown in U.S. Pat. No. 5,352,861 to Steigerwald et al.
However, steady state thrusters may have several disadvantages. For instance, steady state thrusters may respond sluggishly to changes in their operational status. Steady state thrusters usually require several milliseconds for activation, and then several minutes to reach thermal equilibrium. Moreover, steady state electric thrusters are not ideal for applications requiring only a small thrust or short-duration thrust, because at power levels below a few hundred watts steady state thrusters are commonly unstable and inefficient.
The pulsed thruster applies a series of electric current pulses of limited duration (typically on the order of microseconds to milliseconds, with microseconds being common for the low energy thrusters under consideration here) to the propellant to energize the propellant. A sample schematic of a conventional pulsed thruster system 20 is shown in FIG. 1. The system 20 includes a low DC voltage primary power supply 22, a high DC voltage thruster power supply 24, a control circuit 26, an ignition circuit 28, an ignition device 30, a capacitor 32, and a thruster 34. The primary power supply 22 is coupled to the thruster power supply 24, which in turn is coupled to the ignition circuit 28 and selectively coupled to the capacitor 32. The ignition circuit 28 is coupled to the ignition device 30, such as a spark plug, and receives commands from the control circuit 26. The capacitor 32 is selectively coupleable across the thruster 34.
In operation, the primary power supply 22 provides power to the thruster power supply 24, which charges the capacitor 32. The capacitor 32, in turn, applies this voltage across the thruster 34, which has first and second spaced electrodes 38, 40. In accordance with a signal received from the control circuit 26, the ignition circuit 28 fires the ignition device 30. The firing of the ignition device 30 provides a sufficient amount of energy to cause an arc to form on the surface of the propellant 42 between the first and second electrodes 38, 40, thus completing the circuit with the capacitor 32.
The propellant 42 is introduced into the space 44 between the first and second electrodes 38, 40. The energy released from the arc formed between the first and second electrodes 38, 40 may cause the propellant 42 to change into a gaseous form, and particularly an ionized gaseous form known as plasma. The plasma exits the space 44 at high velocity to provide thrust. As the propellant 42 is heated, the propellant 42, which is in a solid or semi-solid form as shown, is advanced into the space 44 through the action of the force F.sub.S, which represents the force provided by a spring (not shown) which abuts the surface of the propellant 42 to urge the propellant 42 into the space 44.
Pulsed thrusters have several advantages compared to steady state thrusters. For example, the time required to activate a pulsed thruster is generally shorter than for a steady state thruster. Pulsed thrusters may achieve thrust in a short time duration, typically microseconds, compared to the time in which a steady state thruster can be turned on and off, typically seconds. Pulsed thrusters also generally achieve a higher peak power level, resulting in high momentum impulses compared to steady state thrusters. Also pulsed thrusters can easily vary their average thrust level by varying the capacitor energy and the pulse rate (pulses per second). Further, the pulsed thruster is generally not unstable in lower power applications.
Nonetheless, pulsed thrusters have their disadvantages. For instance, the circuit elements used to provide the electrical discharge may be subjected to high stresses, and consequently may have a relatively short useful life.
Additionally, current ringing or oscillation can occur in the capacitor and the thruster. Ringing occurs when current continues to flow back and forth through the circuit after the initial discharge of the capacitor, energizing inductances in the lines connecting the capacitor 32 with the thruster 34. FIG. 2 shows a plot of two consecutive current oscillations (A and B) in the capacitor 32 associated with current pulse discharges at times t1 and t2, respectively for the circuit of FIG. 1. The vertical axis represents current level and the horizontal axis represents time, and a typical pulse length T.sub.C is illustrated.
Ringing can cause damage to the entire system 20. For example, ringing may result in the charging of the capacitor 32 against its normal polarity, which may increase the wear on the capacitor 32. Additionally, current reversal through the capacitor 32 can result in considerable energy loss, which degrades overall thruster efficiency and also increases capacitor wear. Further, the corresponding current oscillations through the thruster 34 tend to increase heating of the conductors 46, 48 which connect the capacitor 32 to the electrodes 38, 40 within the thruster 34 and to increase heating of the electrodes 38, 40, the thruster insulators (not shown) and the propellant 42. This increased heating tends to produce undesirable erosion of the electrodes 38, 40 and insulators within the thruster 34, potentially shortening their life. Further, ringing can result in reversal of thrust forces within the thruster, reducing both thrust and efficiency.
It has been suggested that the ringing in the system 20 may be reduced by coupling a diode in parallel with the capacitor 32 and the thruster 34. Specifically, such a solution is suggested by Kimura et al. in Preliminary Experiment on Pulsed Plasma Thrusters with Applied Magnetic Fields, presented at the 13th International Electric Propulsion Conference (1978). In particular, Kimura et al. suggest that the diode in parallel with the capacitor and the electrodes of the thruster may eliminate the oscillatory nature of the main discharge. This solution, however, still allows some undesirable reversal of current in the system.
Furthermore, in a conventional thruster system, as is shown, the impedance of the thruster 34 is significantly larger than the impedance of the capacitor 32 to ensure that most of the energy is delivered to the thruster 34 when the capacitor 32 discharges. Simply put, the capacitor 32 and the thruster 34 will participate in the energy distribution after the capacitor discharge in proportion to their relative impedances. Given that the capacitor is typically on the order of 10 m.OMEGA., for 80% of the energy to be distributed to the thruster 34, the impedance of the thruster 34 must be on the order of 40 m.OMEGA.. The energy distributed to the capacitor is generally lost through heating of the capacitor 32.
However, increasing the impedance of the thruster 34 decreases the efficiency of the thrust production in the thruster 34. For a thruster 34 relying on electrothermal effects (the production of flat through creation of high pressure), increases in thruster impedance can result in excessive propellant ablation and reduced thruster exhaust velocity. For a thruster 34 relying on electromagnetic effects (the production of thrust through electromagnetic forces), increases in thruster impedance can also result in decreased thrust per pulse. For a thruster 34 relying on both electrothermal effects and electromagnetic effects, the effects may be cumulative.